As is known, one typical type of helicopter has a main rotor which produces lift and forward thrust in response to torque provided thereto through rotary power means, including an engine, in a manner determined by the pitch angle of the rotor blades. In order to stabilize the airframe against rotation as a consequence of the torque applied by the engine to the main rotor, this type of helicopter has a tail rotor, the rotary speed of which is a fixed geared function of the speed of the main rotor. The angle of the tail rotor blades, in addition to being adjustable to provide maneuvers in yaw, is adjusted to provide thrust to apply torque to the airframe, about a yaw axis, which will compensate for the torque applied by the engine to the main rotor, so that the airframe will be rotationally stable rather than tending to rotate under the rotor. One method of control over the amount of compensating thrust provided by the tail rotor is achieved by varying the tail rotor pitch angle as a function of the amount of collective pitch angle of the main rotor blades.
Thus, the tail rotor blade angle pitch beam is provided with a command component which bears a fixed ratio to the collective pitch command to the main rotor, referred to as collective/tail pitch mixing or proportional blade pitch mixing.
The problem with proportional blade pitch mixing of the type described hereinabove is that it is incapable of providing a correct amount of torque compensation for both dynamic and steady-state operating conditions of the aircraft. For example, it is well known that the relationship between main rotor torque and collective pitch is not constant with forward speed increasing from hovering flight. Additionally, because of changes in aerodynamic wake patterns at the tail of the aircraft, the relationship between tail rotor thrust and tail rotor collective pitch may vary. Additional variability in the desired tail rotor blade pitch angle compared with main rotor collective pitch is introduced with variations in helicopter weight, center of gravity, climb speed, altitude and air temperature. Thus, the proportional collective/tail rotor mixing is at best a compromise which must be compensated, either by the pilot repositioning anti-torque pedals or by an automatic heading hold system, even during steady flight conditions.
Similarly, under dynamic conditions such as maneuvering transients, the proportion of collective pitch may bear no relation at all to the torque being imparted by the engine to the main rotor drive train and thus, being reacted on the airframe. For example, when the rotor is declutched from the drive train during auto-rotation, collective pitch stick manipulation will cause direct input of tail rotor blade pitch angle through collective/tail pitch mixing, causing the generation of unwanted yaw moment on the airframe which must be neutralized by the pilot or the heading hold system. Similarly, when resuming power flight following an auto-rotation descent, the engine is initially at idle, but the rotor speed has increased to the level necessary to maintain the desired rate of descent. In order to arrest the descent, a large positive collective pitch input is required. This results in a corresponding yaw moment introduced by the collective/tail mixing. Initially, as long as the rotor speed exceeds the engine speed, the engine is still approximately at idle and there is no airframe torque reaction because the rotor is still disengaged from the engine. Therefore, any torque compensation by collective/tail mixing will cause a yaw moment which must be compensated for by the pilot or heading hold system. As soon as the rotor slows sufficiently that its speed is equal to or below that of the engine, the clutch reengages, initially causing a droop in engine speed due to the torque loading thereof; the engine speed in attempting to regain the reference engine speed, spools up at a high rate and provides a torque overshoot. At this stage of the transient, there is inadequate torque compensation so that the pilot must reverse his pedal input. These examples illustrate that under both steady-state and maneuvering flight conditions, main rotor collective pitch is an inadequate indicator of the level of engine torque reaction on the airframe which must be compensated by the tail rotor.
In response to the problems associated with proportional blade pitch mixing of the type described above, U.S. Pat. No. 4,493,465 to Howlett et. al. provides an improved means for stabilizing the helicopter airframe against rotation which adjusts the blade pitch angle of the helicopter tail rotor in response to a parameter indicative of gas generator speed in a helicopter having a free turbine engine. The parameter may be a direct measurement of gas generator speed, or other suitable indicator of engine speed, e.g., fuel supply, free turbine or other output shaft torque or acceleration. Said patent greatly improves the torque compensation provide by the tail rotor by controlling tail rotor pitch with a signal which more accurately represents main rotor induced torque on the airframe.
However, it has been found that the torque compensation provided by said patents is not correct for various dynamic and steady-state conditions of the aircraft. For example, a parameter indicative of gas generator speed is indicative of total engine torque, rather than main rotor torque, and does not account for that portion of main rotor torque required to drive the tail rotor and helicopter auxiliary equipment, e.g., generators and hydraulic pumps. Therefore, compensation based on the gas generator speed signal alone may produce a slight yaw transient which must be compensated for either by the pilot or by the heading hold system.
Additionally, the patent to Howlett et. al. neglects the torque contribution generating by the aerodynamic forces on the airframe, e.g., the aerodynamic forces on the fuselage and tail surfaces. Additionally, in newer generation helicopters, a shrouded fan rather than a conventional tail rotor may be used for torque compensation, and aerodynamic forces on the shroud affect both the torque on the airframe and the performance of the tail fan. As used herein, the term "tail rotor" is intended to refer to helicopter anti-torque devices including both the convention type of torque compensation device, i.e., a tail rotor, and the newer generation tail fan devices. The aerodynamic forces on the tail portion of the aircraft include those forces on the tail fan shroud of a helicopter using a tail fan as a torque compensation device.
At high speeds, said patent may provide compensation that would not be required or desired. For example, at 120 knots, one type of helicopter will generate sufficient aerodynamic force not to require any contribution from the tail rotor whatsoever. Therefore, the tail rotor contribution may tend to cause slight yaw variations or biases in the heading of the aircraft. These biases remain almost invisible to the pilot because of compensation by a heading hold or stability augmentation system (SAS). However, a torque compensation system which does not take into account aerodynamic effects and allows the SAS to drive the final aircraft heading solution will not result in as accurate pointing of the vehicle, which will decrease the effectiveness of a ballistic solution. For example, even if the pilot cannot tell the difference between a 0.25 degree amplitude oscillation in the yaw axis and a 0.1 degree amplitude oscillation, the latter solution will increase the kill probability because of reduced projectile scattering.